Autonomous precision navigation

ABSTRACT

The new invention is a method for a self-contained multifunctional navigation device. It computes 3D-Spatial velocity of the vehicle and of fluid medium surrounding the vehicle, using a movement sensor system comprising low power Transmitter and plurality of Receivers placed close to the Transmitter. The said computation measures propagation time delay of low frequency pulse envelope modulating carrier EM (radio) waves using Time to Digital Converter (TDC). Orientation, direction, position and movement data are derived using well known mathematical formulae. Derived movement data are displayed graphically on a Visual Display Screen. A commercial computer comprising real time operating system, stored digitalized Navigation data and digitalized map facilitates data computation and control and guidance. The method and movement sensor device can be used in any type of vehicle (airborne, surface, sub-surface, marine, sub-marine or space) as a means for navigation and control guidance aid.

FIELD OF INVENTION

This invention relates to a method and an on board self-containedmultifunctional system for Navigation and Control of any type of vehicle(airborne, surface, sub-surface, marine, sub-marine or space) on whichthe system of this invention is mounted. In particular it relates tomeasuring spatial velocity along body fixed axes of the vehicle on whichthe system is mounted by measuring propagation time delay for apredetermined time frame pulse that modulates carrier wave & travelsthrough space from a generating source (Transmitter) to multiplereceivers each of which is placed on one of the principal axes(longitudinal, lateral and vertical) of the subject vehicle. Movementand orientation data with respect to geophysical plane are derived usingthe said spatial velocity and well known mathematical formulae. Theinstantaneous position, course and altitude are derived by deadreckoning (DR) from an initial setting. For vehicles moving throughfluid (air and/or water), additional receivers placed at an offset angleto receivers placed along longitudinal, lateral and vertical axes of thesubject vehicle facilitates computation of dynamic data of fluid (airand/or water) in the vicinity of the subject vehicle moving in the saidmedium(s) and effect of movement of the said medium(s) on the movementof said vehicle. Using a commercial computer and stored digitalized mapand digitalized navigation data and display unit a multifunctional aidfor Navigation and Control of the said subject vehicle is provided. Themethod and system can serve both manned and unmanned vehicles.

BACKGROUND OF THE INVENTION

Navigation aids are meant to determine speed, direction, orientation,position of a vehicle and these data are used to control and guide thevehicle for safe passage from start to destination. For vehicles movingin fluid medium(s) [air and/or water], the relative speed and directionof the medium in the proximity of the vehicle and effect of movement ofmedium on the vehicle are also needed to be determined.

For these functions, multiple aids were being used. But rise in trafficdensity, increase in speed, widening of area of operation, emphasis oneconomic and fast efficient transit have shifted the focus to new typesof Navigation aids which are multifunctional, more accurate andreliable.

At present, such aids can be put under two categories. Satellite basedNavigation System GNSS (Global Navigation Satellite System) and INS(Inertial Navigation System). In some systems both the systems arecombined to overcome the short comings of both.

GNSS comprising of GPS (USA), GLONASS (RUSSIA) satellites along withaugmentation systems DGPS, WAAS (USA), EGNOS (Europe), MTSAT (Japan),GAGAN (India), is being groomed to evolve as primary/sole means ofNavigation. Precision (accuracy) of satellite based Navigation is wayahead in comparison to other Radio Navigation Aids and is in the rangeof 3-30 metres for civilian use. It has the global coverage except inextreme polar-regions. But it has not the other basic performancerequirements: Continuity, Availability and Integrity up to the expectedlevel as yet. Satellite based Navigation Systems also suffer from errorsthat can degrade its precision. The errors are—ionospheric, atmospheric,clock, receiver, ephemeris and position dilution of precision. And ofcourse, the satellite constellation owner country can degrade theprecision and even suspend availability of service. Possibilities ofblack out and jamming are other shortcomings of this type of Navigationaids.

INS is the mainstay of space, marine, submarine and Guided Weapon GN&C(Guidance, Navigation and Control). But it plays supplementary role forcommercial vehicle Navigation of other types of vehicles. This is due to(i) high cost of equipment, (ii) mechanical parts hampering reliabilityin the form of wear, gimbals lock and ring lock (laser) and puttingrestriction on maneuver, (iii) requirement of position correction byinput from other type of Navigation Systems during operation period toovercome integration drift.

Both the systems are unable to determine fluid (air and/or water)movement data in the vicinity of the vehicle through which the vehiclemoves and effect of said fluid movement on vehicle movement which is avery important determinant for movement of vehicle through fluid (airand/or water).

The shortcomings of Satellite based Navigation Systems and InertialNavigation Systems are overcome by method and system of this newinvention.

OBJECTS OF INVENTION

The primary object of this invention is to provide a self containedonboard multifunctional Navigation and control guidance System toeliminate external influences inducing error and reducing dependence onmultiple external aids.

It is another object to provide a Navigation system which fulfils allthe four basic performance requirements i.e. a Navigation aid which willprovide (i) uninterrupted service during complete operation period(Continuity), (ii) maintain required quality all the time(Availability), (iii) will warn the operator/crew in the event offailure (Integrity) and (iv) provide guidance within predefined (such asconformity to Required Navigation Performance (RNP) for air navigation)tolerance (Precision).

An important object is to compute spatial velocity of the vehicle alongbody-fixed axes (longitudinal, lateral and vertical) of millimetre(sub-centimetre) accuracy.

Another object is to determine speed, direction, orientation of thevehicle with respect to surface of the earth and for vehicles movingthrough fluid medium speed of the vehicle relative to fluid medium.

An important object is to provide displacement magnitude, rate ofdisplacement and instantaneous geo-physical position of the vehicle withprovision for reliable means for self-correction for drift integrationerror.

Another object is to compute relative or absolute speed and direction ofthe fluid medium (air and/or water) movement in the vicinity of thevehicle through which the vehicle is moving and effect of fluid mediummovement on the vehicle electronically without pressure operatedinstruments and without any part or portion of the system placed on theexternal skin of the vehicle.

Another object is to provide these basic data and data derived fromthese basic data to aid safe, efficient, economic operation of all typesof vehicles (air, surface, marine, submarine and space) and providingguidance and control to the vehicle from start to destination.

SUMMARY OF THE INVENTION

Autonomous Functional Aspects:

By Autonomous Functions is meant the functions the system of presentinvention can perform in complete isolation from any instrument orsystem whether located onboard or outside of the user vehicle.

For user vehicles of all categories the common autonomous functionalaspects are as follows:

-   i) One aspect of this invention is the system of this invention has    a self-contained sensor for sensing spatial velocity of the subject    vehicle along its longitudinal, lateral and vertical axes.-   ii) Another aspect is determination of momentary position as well as    orientation (yaw, roll, pitch angles) with respect to earth's    surface and movement data of the mount vehicle under all operating    condition and anywhere around the globe continuously in complete    isolation from any instrument or system external to the system of    invention, using the self-contained sensor output.-   iii) Another aspect is, the sensor can compute relative movement    data of fluid (air and/or water) in the proximity of the user    vehicle moving through the said fluid and compute the effect of the    said fluid movement on the vehicle movement continuously.-   iv) The system can measure any range of velocity in either direction    (forward or reverse) along each of the spatial axis.

In respect of the preferred embodiment for use in air vehicles anddiscussed in this disclosure, the autonomous functional aspects are asfollows:

-   i) The system in its embodiment for use in air vehicles can    substitute all the following conventional primary flight instruments    or PFD's (in recent use) of air vehicles i.e. Altimeter, Airspeed    Indicator, Attitude Indicator (Artificial Horizon), Magnetic    Compass, Heading Indicator/Horizontal Situation Indicator, Turn    Indicator, Vertical Speed Indicator and additional Panel instruments    i.e. Course Deviation Indicator, Radio Magnetic Indicator and    graphically display the relevant information provided by the said    panel instruments on a single visual display unit numerically and/or    non-numerically.-   ii) Both course on ground (track) and mount vehicle heading (along    its longitudinal axis) will be provided simultaneously and for both,    heading indications with respect to magnetic and true north will be    graphically displayed as numerical reading in degrees.-   iii) In addition the information regarding vital speeds will be    indicated alongside TAS (True Airspeed) graphical display. Wind    direction and speed, wind vector components, angle of attack and    drift rate will also be displayed.-   iv) Another aspect of the present invention is it will provide    instantaneous position of the mount vehicle by displaying position    of the subject vehicle on moving map display as well as in    Coordinates (Longitude/latitude) based on WGS-84 datum reference in    a separate window. The position information is also displayed in    respect of relevant way points on the said display unit.-   v) Another aspect of this invention is to provide Navigation and    Control Guidance by providing a path (to be followed as predefined)    overlay on the said map display and can generate offset path overlay    too.-   vi) One more aspect is that, the system will sense initiation of    drift instantaneously and help initiation of immediate corrective    action in real time so that drift is kept to minimum or negligible.    Further, the system can measure distance deviated from track and    course correction required to resume the track instantaneously. To    facilitate control movement for track alignment a group of    information is displayed. The group information comprises ground    speed, ground heading (track), off track distance, track angle    error, set course to, turn anticipation (with turn radius    calculation and determination of position of origin of turn radius),    indication for commencement and discontinuance of turn and    climb/descent instructions and a non-numeric graphic display with an    icon as aim figure for track alignment will make it easier for the    crew to maintain track.-   vii) The system of the present invention will generate a new path    overlay on map as and when deviation from predefined primary path is    imminent and an alternate destination point is selected due to any    emergency or any other reason or due to selection of an offset    track.-   viii) An important aspect is that the system of present invention    can generate a flight plan and edit the same during flight without    affecting normal operation.-   ix) The system will compute distance and bearing to any way point    from aircraft position or from any other way point and display the    same.-   x) The system will compute slant range and bearing to any way point    and display the same.-   xi) The system enhances the easiness of controllability of the    subject vehicle by providing displayed direction for speed    adjustment required, altitude (climb/descend to), turn (to heading)    directions and can feed. Automatic Flight Control System (auto    pilot) with data required for controlling the vehicle.-   xii) The present invention will provide Navigation and control    guidance for VFR/IFR flight rules as required.-   xiii) Another important aspect of the present invention is to    provide Approach and Landing Guidance of Approach Category I, II,    III and virtual ground controlled approach (audio guided) of    Precision Approach Radar (PAR) type to aircraft for the runways    around the world for which longitude, latitude (coordinates based on    WGS-84 datum) Information is available.-   xiv) The system of this invention can cater for any shift in runway    touchdown point and change in runway usable length.-   xv) The system of this invention will also provide Navigation and    Control guidance on the ground during taxi out and taxi in phases of    movement.-   xvi) The system of present invention is not dependent on any    barometric or inertial navigation instrument or earth referenced    navigation instrument.-   xvii) The system also provides the 4^(th) Dimensional i.e. solutions    to Time related problems.-   xviii) The system will comprise of a commercial computer    encompassing Real Time Operating System (RTOS) for real time    computation and derivation of related data.-   xix) The system will consist of a storage memory in which    digitalized maps, charts, navigation data, geo-physical data check    lists and procedures will be stored.-   xx) Provision of backup power supply unit will enable the system of    this invention to provide all the services of autonomous function    enumerated above even in the event of power supply failure from main    source of supply. Thus the subject vehicle will continue to have    service of the system and reach a destination safely.

Integrated Functional Aspects:

The system of invention, when integrated with equipments on board thesubject vehicle, can perform the following functions too.

-   i) Integrated with Automatic Flight Control System (AFCS) commonly    known as auto pilot, the system of this invention can have dynamic    control over pitch, roll and yaw and thrust commands. Since the    controlling inputs are electronically generated the problem of    violent response as happens in the event of vacuum operated    instrument failure is avoided.-   ii) Integrated with digitalized data transmission/reception of Air    Traffic Network (ATN) or suitable communication means, the system of    present invention will enhance the Automatic Dependent Surveillance    (ADS) capability of Ground Controlling Agencies by providing    position and other required data as laid down by ICAO for required    surveillance performance (RSP) continuously.-   iii) Digitalised data transfer between aircraft directly (ADS-B) or    through ground establishment, will provide full proof collision    avoidance on ground or in the air because of (i) high measurement    accuracy in millimetres (linear displacement) and seconds (angular    displacement) per second (time) computed in real time, (ii)    computation of collision probability and point of prospective    collision (resolved 3-dimensionally) which will enable the crew to    take procedural avoidance method well in time to prevent collision.    Display of surrounding traffic on moving map display (CDTI) will    enhance traffic situational awareness of the crew.

EMBODIMENTS

The system of new invention can be used in any type of vehicle (air,surface, marine, sub-marine or space). The sensor sub-system operationbeing same only computed Navigation Aid data and Control data will varyto cater for user specific requirement. Hence the embodiments for use indifferent types of vehicles will substantially be same.

The Preferred Embodiment

The acronym for Autonomous Precision Navigation is ‘APNA’ which, inIndian. National Language (Hindi) means ‘my own’. The method of thisinvention provides an onboard system which is truly ‘my own’ NavigationAid for the mount vehicles. To understand the principle and applicationof the invention, embodiment for use in aircraft which the inventorcalls ABLAN (Airborne Landing and Navigation) System will be explained.This embodiment is illustrated in the drawings and specific languagewill be used to describe the same.

Nevertheless, may it be understood that the scope of the invention isequally applicable to other classes of vehicles (surface, sub-surface,marine, submarine or space) too for solving navigation, control andsafety related problems of those classes of vehicles as it wouldnormally occur to a person skilled in the art to which the inventionrelates.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 Shows a hardware architecture of the system.

FIG. 2 A Depicts the principle of movement sensor (static condition)

FIG. 2 B Depicts the principle of movement sensor (dynamic condition)

FIG. 3 A Shows the body fixed axes along longitudinal, lateral andvertical directions of the mount vehicle.

FIG. 3 B Shows the dynamic Geophysical axes in three dimension.

FIG. 3 C Shows the combination of body fixed axes of FIG. 3 (a), anddynamic Geophysical axes of 3 B.

FIG. 4 A Shows primary movement sensor Antennae layout along threespatial axes.

FIG. 4 B Shows additional movement sensor Antennae for computation ofwind direction, speed and component vectors.

FIG. 5 Shows hardware architecture for movement sensor data computation.

FIG. 6 Shows transformation of spatial movement data (forward) toGeophysical movement data (forward and vertical) under without bankcondition.

FIG. 7 Shows the same transformation of FIG. 6 under bank condition anddetermination of pitch angle and bank angle.

FIG. 8 A Shows method for calculation of horizontal movement vectoralong dynamic geophysical (OX) axis.

FIG. 8 B Shows method for calculation of speed along vertical (OZ) axis.

FIG. 8 C Shows method for calculation of speed along lateral (OY) axis.

FIG. 8 D Shows method for calculation of yaw Angle and ground speed.

FIG. 9 A Shows method for calculation of turn anticipation and radius(fly by).

FIG. 9 B Shows method for calculation of turn anticipation and radius(fly over).

FIG. 10 Shows the Avionics section of display.

FIG. 11 Shows display of information of all sections except AvionicsSection.

FIG. 12 Depicts the single screen display of primary flight data,navigation and control guidance and other information.

FIG. 13 A Shows method for computation of visual display for LNAV duringapproach.

FIG. 13 B Shows method for, computing visual display for VNAV duringapproach.

FIG. 14 A Shows method for computing audio guidance for LNAV duringapproach.

FIG. 14 B Shows method for computing audio guidance for VNAV duringapproach.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 depicts the hardware structure of the system of the newinvention. It has a sensor sub-system (4) comprising transmitter andreceivers for sensing 3-dimensional velocity of the mount vehicle and offluid (air and/or water) in the vicinity and processor for computing thesaid data. The outputs of the sensor are fed to the CPU (5) of acommercial computer. The said computer encompasses real time operatingsystem (RTOS) and high level language. It derives the instantaneousmovement data and orientation of the mount vehicle using spatialvelocity derived by sensor sub-system and instantaneous positioninformation by dead reckoning using derived displacement information.Digitalized map data and navigation data stored in the memory,facilitate control guidance, flight plan generation and compute terrainclearance. The derived information is graphically displayed on a singlescreen (1). The crew has option to use key-board (7) to use the computerfor retrieving stored information/chart/checklist etc. for readyreference and assembling/editing flight plan. Provision of back-up powersupply (6) ensures uninterrupted operation of the system even in theevent of failure of main power source. The sub-sections of the systemdiscussed so far are meant for stand-alone autonomous function.Integration with Automatic Flight Control System (AFCS) (2) enables thesystem of present invention to have dynamic control over the subjectaircraft. Integration with Communication Network (3) for digitalizeddata transfer with traffic elements and ground stations for Air TrafficManagement (ATM) will enhance Automatic Dependent Surveillance (ADS),traffic flow management capability and ensure full-proof operationalsafety (ACAS-I &II) on the ground and in the air.

Principle

As shown in FIG. 2 A, EM (radio) wave travelling from a transmitterantenna (8) to a receiver antenna (9), both mounted and fixed on thevehicle and aligned to its principal axes (longitudinal, lateral orvertical) have the constant separating distance (S_(o)) (10) and duringmovement their spatial velocity in the direction of the principal axeswill be same as that of the mount vehicle. EM wave has absolute velocity(V_(EM)) and is equal to 3×10⁵ Km/Sec. i.e. 3×10¹⁰ Cm/Sec (Theapproximate value is taken for explaining the invention. However theactual value of 299999.998 Km/Hr will be used for computation of data inactual system). Every EM wave will take a constant time (t_(o)) (11) totravel through space from Tx. Ae. (8) to Rx. Ae (9) when the vehicle isstatic.

As shown in FIG. 2 B, when the mount moves at a velocity of V_(M)Km/Hour in the direction of Rx. Ae., even though (S_(o))(10) remainsunchanged, a wave travelling from Tx. Ae. (8) to Rx. Ae. (9) will haveto traverse an additional distance (ΔS)(13) which is the distancetravelled by the mount in (t_(o)) (11) Sec. Now, instead of (t_(o)) (11)Sec the time taken by EM wave to travel from Tx. Ae. (8) to Rx. Ae. (9)will increase. Let the additional time be (Δt) (14) Sec. Similarly,movement in opposite direction will cause reduction of distance by (ΔS)and time taken by (Δt) Sec. Thus, spatial velocity of the mount V_(M) isreflected in the Rx. Ae. (9)

The fact will give rise to a difference in number of waves (cycles)received during movement as compared to number of waves that would havebeen received if the mount were in static condition in a certain timeframe. Let this difference in no. of cycles in the given time frame beΔƒ. Δƒ is a dependent function of Vm. Hence, mount velocity alongspatial axes can be computed when Δƒ is determined.

For the purpose of compution, we will relate Δƒ=1 cycle to V_(M)=1Km/Hour. (It may be noted that we can adopt whatever conversion ratio isconvenient for the mount in consideration and unit of measurement canalso be chosen to meet the user's requirement).

Basic Equations

Let, distance between Tx. Ae. and Rx. Ae.=S_(o) Cm.Time to traverse (S_(o)) Cm by EM wave=t_(o) Sec.Frequency of EM wave=f Hz.

Mount Velocity (V_(M))=1 Km/hour

$\begin{matrix}{t_{o} = {\frac{S_{o}}{V_{EM}} = {\frac{S_{o}}{3 \times 10^{10}} = {\frac{S_{o}}{3} \times 10^{- 10}{{Sec}.}}}}} & (1) \\\begin{matrix}{{\Delta \; S} = {{to} \times {V_{M}\left( {{in}\mspace{14mu} {Cm}\text{/}{Sec}} \right)}}} \\{= {\frac{S_{o}}{3} \times 10^{- 10} \times \frac{10^{5}}{3600}{Cm}}} \\{= {\frac{S_{o} \times 10^{- 7}}{108}{{Cm}.}}}\end{matrix} & (2) \\\begin{matrix}{{\Delta \; t} = \frac{\Delta \; S}{V_{EM}}} \\{= \frac{S_{o} \times 10^{- 7}}{108 \times 3 \times 10^{10}}} \\{{= {\frac{S_{o}}{324} \times 10^{- 17}{{Sec}.\mspace{14mu} \left( {{for}\mspace{14mu} {one}\mspace{14mu} {cycle}} \right)}}}\;}\end{matrix} & (3) \\{{{For}\mspace{14mu} {frequency}\mspace{14mu} f\mspace{14mu} {Hz}},{{{duration}\mspace{14mu} {of}\mspace{14mu} {one}\mspace{14mu} {{cycle}\left( {p.d} \right)}} = {\frac{1}{f}\mspace{11mu} {{Sec}.}}}} & (4)\end{matrix}$

The cumulative delay ΣΔt to be equal to 1/ƒ Sec. so that Δƒ=1 cycle, weare to cumulate time delay for

$\begin{matrix}\left( \begin{matrix}{{\left. {{\frac{1}{f} \div \Delta}\; t} \right){Cycles}} = {\frac{1}{f \times \Delta \; t} = {\frac{324 \times 10^{17}}{{So} \times f}{Cycles}}}} & \;\end{matrix} \right. & (5)\end{matrix}$

ΣΔt for no. of cycles in equation (5) based on V_(M)=1 km/hr is therequired time-frame (t_(c)) where,

$\begin{matrix}{t_{c} = {\frac{324 \times 10^{17}}{{So} \times f \times f} = {\frac{324 \times 10^{17}}{{So} \times f^{2}}{{Sec}.}}}} & (6)\end{matrix}$

Equation (6) is the duration (t_(c)) for which it is required to countthe received cycles to relate Δƒ (cycles):

$V_{M{(\frac{Km}{Hour})}}\text{:}$

1:1 i.e., Number of cycles directly reads spatial velocity (in Km/Hour)along the axis in consideration.

When measured from time 0 (zero) of time frame (t_(c)), for movement inthe direction of the Rx. Ae. or opposite to it in the time frame ofequation (6), number of pulses received will be equation (5) ±n wheren=V_(M) in Km/Hour. The value of (n) is (+) positive for movement in thedirection of the receiver antenna and (−) negative for movement in theopposite direction. However, the method requires an absolutely stablecarrier frequency because even a single cycle deviation will not beacceptable and as is well known in the art such standard is not possibleto achieve. To overcome the difficulty a new approach is introduced inthis invention. Frequency stability may not be achievable but a pulse ofrepresentative duration (t_(c)) can be made stable. A pulse of duration

$t_{c} = {\frac{324 \times 10^{17}}{{So} \times f^{2}}{{Sec}.}}$

will represent the time frame for which the carrier waves be modulatedand instead of counting cycles the propagation, delay of the time framewill be measured accurately by time to digital converter (TDC) of theprocessor of the sensor. The ΣΔt thus computed when divided by pulseduration of one cycle of the predetermined frequency ‘f’ based on whicht_(c) is determined, the quotient will directly read velocity inkilometre per hour along longitudinal, lateral or vertical axis of thevehicle. The other movement data like orientation, displacement, rate ofdisplacement, instantaneous position and many more in respect of thevehicle can be derived using thus computed spatial velocity of thevehicle as has been shown in this disclosure.

Spatial 3-D Velocity of Mount

In the preferred embodiment for disclosure of this invention which is tobe used in aircraft, the system function is enumerated herein.

As shown in FIG. 3A aircraft like any other vehicle has three principalbody fixed axes, viz. longitudinal ({right arrow over (OA)}) (15),lateral ({right arrow over (OB)}) (16), and vertical ({right arrow over(OC)}) (17), and they are spatial in nature. These axes are differentfrom axes oriented on Geo-physical plane which have dynamic co-ordinatesalong {right arrow over (OX)}, (18) {right arrow over (OY)} (19) onhorizontal plane and {right arrow over (OZ)} on vertical plane as shownin FIG. 3B. The word dynamic denotes that the geophysical axes are notfixed to any direction rather they are projection of spatial axes of thevehicle on earth's surface such that X and Y co-ordinates move inalignment with corresponding body fixed axes and Z co-ordinate isorthogonal to the horizontal geo-physical co-ordinates. The origin ofthese two frames is same as shown in FIG. 3C.

As shown in FIG. 4A an omni-directional Tx. Ae. (22) is placed at theorigin and on each of the three principal axes of the vehicle or onlines parallel to principal axes and at an angle 90° to each other a Rx.Ae. (23, 24, 25) is placed at a distance (S_(o)) from the Tx. Ae. (22).

As shown in FIG. 5, a crystal controlled local oscillator (33) feeds theTx. Ae. (22) through a modulator/driver (34) for a duration of equation(6).

Since all the sensor processing channels are identical, only function ofone channel (as shown in FIG. 5) will be explained.

The very low power transmission of EM (radio) wave is received atcorresponding Rx. Ae. (23, 24 or 25). Receiver (37) receives the timeframe pulse the propagation time delay of which is proportional tospatial velocity of the mount along the axis. The received pulse isamplified (37), detected (38) and compared to the original modulating,pulse. The time difference between trailing edges of the two pulses ismeasured using Time to Digital Converter (TDC) (39). This is the timedelay ΣΔt which when divided by 1/ƒ sec gives mount velocity inkilometers per hour along the axes on which Rx Ae is mounted. SpatialVelocity processor (40) carries out this processing and gives output(41). This output is fed to CPU of computer which carries outmathematical calculations to compute data for aiding navigation andcontrol of the vehicle and generate numeric/non-numeric graphic fordisplay of information on VDU.

Let, the sensor output (spatial velocity SV) be along

Longitudinal axis - - - SV_(ƒ)Lateral axis - - - SV_(l)Vertical axis - - - SV_(v)

These outputs of sensor sub-system are spatial velocities of the mountvehicle along respective body fixed axis and gives direct numericalvalues in Km/Hour.

Clock pulse generated by ck pulse gen (35) synchronises the operation ofTx, Rx and processor and other sub-systems. Modulator pulse generator(36) generates main modulating time frame pulse and test pulse.

It may be noted that for the purpose of computation of spatialvelocities in 3-Dimension, the sensors may be a pair of Tx-Rx usingradio, laser or any such travelling EM waves. The basic computationprinciple will stand good. The preferred carrier for this invention isradio wave as the hardware for the system will be cheap and the actualcarrier frequency being irrelevant for the purpose of computation ofvelocity, a suitable allotted navigation aid frequency can be used.

The unit of measurement can be converted to user specific unit e.g.knots and feet per minute which is the case for embodiment for use inaircraft.

Transformation of Spatial Velocity to Velocity Along Geo-Physical(Horizontal & Vertical) Co-Ordinates:

FIGS. 6 & 7 shows method for determining 3-dimensional velocity alonggeo-physical axes and also method for computing yaw, roll and pitchangles.

Derivation of required information is obtained by using well knownMathematical formulae. Trigonometric relations have been used forcomputation for explaining this invention. However, as it would normallyoccur to a person skilled in the art that the problems in issue can besolved by using vector algebraic method, 3-D co-ordinate geometry andmatrix algebra as well.

Determination of Horizontal and Vertical Component of SV_(ƒ) and PitchAngle.

As shown in FIG. 6, Let {right arrow over (OA)} (15) representing SV_(ƒ)be inclined to {right arrow over (OX)} (18) the X-coordinate ofgeophysical plane by θ₁° (21), where θ₁° is the pitch angle.

-   -   Let {right arrow over (OF)} (48) (=F) be the horizontal        component of {right arrow over (SV_(ƒ))} along {right arrow over        (OX)} (18), F being velocity along {right arrow over (OX)} (18)        on Geo-physical plane.

Hence, F={right arrow over ((SV_(ƒ)))}Cos θ₁  (7)

-   -   Let, {right arrow over (OE)} (49) (=V) be vertical component of        {right arrow over (SV_(ƒ))} along {right arrow over (OZ)} (20),        V being velocity component along {right arrow over (OZ)} (20)        the Z-coordinate of geophysical plane.

Hence, V={right arrow over ((SV_(ƒ)))}Sin θ₁  (8)

Since, sensor Rx Ae along {right arrow over (OC)} (17) of aircraft isalso tilted from {right arrow over (OZ)} (20) by θ₁° (21) (Pitch angle)on the same plane, {right arrow over (SV)}_(v) is a projection of V′(eqn.-8) and

{right arrow over (SV)}_(v)={right arrow over ((SV_(ƒ)))}Sin θ₁ Cosθ₁  (9)

As shown in FIG. 7, on bank of angle θ₂° (50), {right arrow over(SV)}_(v) becomes a projection of equation (9) on a plane orthogonal toplane AB and in that position

{right arrow over (SV)}_(v)={right arrow over ((SV_(ƒ)))}Sin θ₁ Cos θ₁Cos θ₂  (10)

From equation-10

$\begin{matrix}{{{Sin}\; \theta_{1}{Cos}\; \theta_{1}} = {\frac{{\overset{\rightarrow}{SV}}_{v}}{\overset{\rightarrow}{{SV}_{f}}{Cos}\; \theta_{2}} = {{> {2\; {Sin}\; \theta_{1}{Cos}\; \theta_{1}}} = {\frac{2\; {\overset{\rightarrow}{SV}}_{v}}{\overset{\rightarrow}{{SV}_{f}}{Cos}\; \theta_{2}} = {{> {{Sin}\; 2\; \theta_{1}}} = {\frac{2{\overset{\rightarrow}{SV}}_{v}}{\overset{\rightarrow}{{SV}_{f}}{Cos}\; \theta_{2}} = {{> \theta_{1}} = {\frac{1}{2}{Sin}^{- 1}\frac{2\; {\overset{\rightarrow}{SV}}_{v}}{\overset{\rightarrow}{{SV}_{f}}{Cos}\; \theta_{2}}}}}}}}}} & (11)\end{matrix}$

Determination of Bank Angle.

As shown in FIG. 7, With bank aircraft attains a lateral movement along{right arrow over (OY)}.

Let, the lateral velocity be L along {right arrow over (OY)}.

This can be resolved into two components along sensors placed on {rightarrow over (OB)} (16) and {right arrow over (OC)} (17) during bank and L(51) makes an angle θ₂ (50) equal to angle of bank with sensor placedalong {right arrow over (OB)} (16) such that component

along {right arrow over (OB)}, SV₁ =L Cos θ₂  (12)

and along {right arrow over (OC)}, SV_(v) =L Sin θ₂  (13)

From equations, (12) and (13),

$\begin{matrix}{{{Angle}\mspace{14mu} {of}\mspace{14mu} {Bank}},{\theta_{2} = {\tan^{- 1}\frac{\overset{\rightarrow}{{SV}_{v}}}{\overset{\rightarrow}{{SV}_{l}}}}}} & (14)\end{matrix}$

Speed on Geophysical Plane.

Resolving the spatial velocities along {right arrow over (OA)}(SV_(ƒ)),{right arrow over (OB)}(SV_(v)) and {right arrow over (OC)}(SV_(l)) intocomponents along {right arrow over (OX)}(F), {right arrow over (OY)}(L),and {right arrow over (OZ)}(V) respectively we get,

F=SV_(ƒ) cos θ₁(FIG. 8 a)  (15)

V=SV_(ƒ) sin θ₁+(SV_(v) Cos θ₁ Cos θ₂−SV_(l) Sin θ₂)(FIG. 8 b)  (16)

L=SV_(v) Sin θ₂+SV_(l) Cos θ₂(FIG. 8 c)  (17)

Accuracy of Velocity Measurement:

ΣΔt=pulse duration (p.d.) of one cycle represents minimum discerniblevelocity (predetermined) which I have taken to be (for the purpose ofexplanation) 1 (one) Km/hour. Also for the purpose of explanationwhatever calculation has been derived are based on Tx Ae. to Rx Ae.distance S₀=3 cm. Use of Time to Digital converter (TDC) of 10Pico-second resolution, which is commercially available, for time delaymeasurement enhances the discernibleness. For example, for referencefrequency 1 GHz, the cumulated time delay of 1 nano-second represents 1Km/Hour velocity and 10 p-sec resolution of TDC will enhance measurementaccuracy to 0.01 Km/Hour i.e., 2.777777778 millimetre per second. For 10GHz reference frequency the enhanced accuracy is 0.1 Km/hour 27.77777778millimetre per second. Availability of TDC of better resolution willfurther enhance the measurement accuracy.

Determination of Position:

Position displacement ‘S’ is obtained by multiplying Pulse RepetitionPeriod (PRP) with Average Ground Velocity (GS_(AV)) for the duration ofPRP. The ‘S’ value is added to previous position in the craft trackdirection to obtain current position. The position information isprovided in longitude and latitude (based on WGS-84 datum) and bydisplay of craft icon on a moving map wherein movement of moving map isproportional to displacement of the craft. The conversion of bearing anddistance information to co-ordinate and vice-versa is obtained by usingVincenty's formula which provides positional accuracy in millimeters.

$\begin{matrix}{{GS}_{AV} = \frac{{GSp} + {GSc}}{2}} & (18)\end{matrix}$GS_(p)=Ground Velocity at the end of previous PRP

GS_(c)=Ground Velocity at the end of current PRP

Position Accuracy:

It is proposed to have two sets of sensor for determination of positionand correction of integration drift error.

The first set is based on 10 GHz reference frequency. This set is meantfor computation of instantaneous 3-D velocity and derivation ofinstantaneous movement data. Determination of current position will becarried out based on the data of this sensor output. At the end of 11second period the position will be corrected by using data provided bythe second set of sensor. The PRP of the first set sensor will be 125millisecond and fix rate is 8 per second. Minimum discernible velocityfor 10 GHz reference frequency is 0.1 km./hr. Hence minimum discernibledisplacement per PRP (125 milli-second) is 3.472222222 millimetre.Maximum Error Probability for delay time measurement based on 10 psecresolution of TDC and 100 psec (pulse duration for one cycle) referenceis 9 psec per PRP. Based on this assumption error probability fordisplacement measurement per PRP is 3.125 millimetre. This will resultin a maximum drift integration error of 90 metres/hour (0.09 Km/hr i.e.,0.048 NM/hr) in the directions along track and cross track. For theembodiment to be used in aircraft, this itself allows for more thantwenty hours of route flying in any airspace conforming to RNP-1 normi.e., containment in ±1.85 Km (±1 NM) envelope.

Self Correction of Positional Error:

As it is intended to use the embodiment for air vehicle fornon-precision approach as well, a second set of sensor is introduced fordrift error self correction. The second layer sensor is based on 1 GHzreference frequency and 10 psec TDC resolution. Its PRP is 11 secs andfor ΣΔt measurement error probability of 9 psec per PRP, displacementmeasurement error is 2.75 cm as against 27.5 cm. drift error for 10 GHzreference frequency over 11 sec operational duration. So position errordrift corrected every 11 seconds will reduce error margin by 10 times.The drift error per hour based on 1 GHz reference frequency is 9 mtrs.

A further correction following the below described method will reducedrift error to almost zero (negligible) magnitude.

Fix rate per hour based on 1 GHz reference frequency is 327. When ΣΔtfor every PRP (11 secs) are integrated for 327 PRPs and compared to sumof TDC measured delay over the same time period, the difference givesthe time measurement error which converted to displacement error willrestrict the drift error to 2.75 cm. The same principle can be practicedfor any operational duration and the equipment error for positiondetermination is a maximum of 2.75 cm. at the beginning of anon-precision approach to any runway.

The above mentioned reference frequencies are given as example and willbe set, keeping operational requirement of various types of vehicles inview.

Integrity Check:

The gap (quiescent period) between transmission of time frame pulses isenough to process the corresponding received pulse as well as to injecta test pulse of predefined duration and the corresponding received pulsefor the test pulse is processed and compared with predefined values atall appropriate levels of processing to check reliability of everysub-system stage. Thus integrity check of the system is ensured prior toevery time frame pulse processing. In the event of failure forpredetermined number of consecutive time frames appropriate failureindication will be displayed and audio warning will be given.

Primary Flight Data

True Air Speed:

The vector sum of value of Spatial Velocity along longitudinal axis ofthe aircraft (SV_(ƒ)) and wind vector along longitudinal axis is TrueAir Speed (TAS) of aircraft in Km/Hour. (Eqn.-29).

Ground Speed:

Ground speed of aircraft is

G.S.=√{square root over (F ² +L ²)} Km/Hour (FIG. 8 D)  (19)

Aircraft Heading/Ground Heading:

ABLAN can deduce both course and track heading by DR (dead reckoning)from an initial setting. The system can also deduce true and magneticheading simultaneously. Both the headings are initialized either atstart up point or line-up point on the runway. The readings are updatedat the end of duration of every time frame pulse and no further helpfrom external source is required after initialization for the rest ofthe operation period.

For updating of ground heading the following method is followed. Valueof SV_(ƒ) sin θ₁ cos θ₁ is compared to that of SV_(v) for wings levelcondition.

If not equal→ground heading and aircraft heading are updated by ±θ₃°.

$\begin{matrix}\left( {{{Fig}.\mspace{14mu} 8}\; D} \right) & \; \\{\theta_{3} = {\tan^{- 1}\frac{L}{F}}} & (20)\end{matrix}$

If found equal→L value is checked.

If L=0, then no drift. Hence, no change in ground or course heading isrequired.

If L≠0, then drift is present. Drift warning in the form of flash oflight under attitude indicator (125 or 127), displacement of aim figurein track alignment indicator (156) proportional to off trackdisplacement (as in FIG. 10), numerical display of off track distance(159), numerical display of track angle error (158), corrected course(required to steer) (160) (as shown in FIG. 11) are provided. Onconfirmation of drift condition, ground heading read out is changed by±θ₃° and aircraft heading read out is not changed.

Second counting time period (PRP) onwards θ₃ of previous counting periodis compared with θ₃ of current counting period.

If θ₃ (Previous)=θ₃ (current), ground heading is not changed.

If θ₃ (Previous)≠θ₃ (current), ground heading is updated by

θ₃ (Previous)˜θ₃ (Current).

Altitude:

Initial setting is done at Runway lineup point. Altitude reading ischecked and if required corrected and set to runway AMSL (data retrievedfrom stored data) altitude. Thereafter the reading is updated at the endof duration of every time frame pulse in accordance with verticaldisplacement per PRP. No external reference is required thereafterduring operation duration at any stage.

Rate of Climb/Descent: (Vertical Speed: VSP)

V value (equation-16) is vertical velocity in Km/Hour. Hence,V×54.68066492 gives the rate of climb/descent in feet per minute.Positive value is indicative of climb and negative value is indicativeof descent.

Angle of Climb/Descent

$\begin{matrix}{\tan^{- 1}\frac{V}{F}\mspace{14mu} {gives}\mspace{14mu} {the}\mspace{14mu} {angle}\mspace{14mu} {of}\mspace{14mu} {climb}\mspace{14mu} {and}\mspace{14mu} {descend}} & (21)\end{matrix}$

(Pitch Angle)

Equation-(11) gives the pitch angle of the craft.

Bank Angle (Roll Angle)

Equation-(14) gives the bank angle.

Rate of Turn (Yaw Angle)

Eqn.-(20) gives the rate of turn.

Turn Anticipation and Turn Coordination:

Turn alert prior to arriving at the way point predefined for enteringinto turn is preset and warning is given by visual as well as audiomeans.

For ‘fly by’ turn (FIG. 9A), the predefined way point to enter turn ispoint (J) on current track, selected ahead of the intersection point (K)of present segment with segment to follow. For ‘flyover’ turn (FIG. 9B),the predefined point is the point of intersection (K) itself. To followthe centre line of the track during turn, a turn radius is determined(if already not notified by ATS route designator). The coordinate of theorigin (O′) of radius is determined. The turn radius (R) to the saidorigin is maintained during the period of turn.

Location of Origin of Turn Radius:

For fly by, the origin (O′) is a point located in a direction of 90° tothe track heading and in the inner side of the turn and is at a distanceR from the entry way point (J) where,

$\begin{matrix}{R = {{Arm} \div {\tan \left( \frac{\phi}{2} \right)}}} & (22)\end{matrix}$

And arm length is the distance between the point to enter turn (J) andthe point of intersection (K) of two segments and φ is the angledifference between present track (JK) and track of the following segment(KL).

For fly over, the origin (O′) is a point orthogonal to the current track(JK) and is at a distance R from the point of intersection (K),

And R=KM sin φ  (23)

Arm length (KL) is the distance between the point of intersection and apoint (L) selected on the track (KL) of the following segment point ‘IA’being mid-point of KL.

The Arm length and turn radius (R) will depend on the type of aircraft.

Wind Vector:

To determine Wind direction, speed and components of wind vector alonglongitudinal, lateral and vertical axes the computation procedure is asfollows.

Craft moving through air will be subjected to two types forces; selfgenerated thrust and wind vector. Along each body fixed axes, theseforces will cause spatial velocity for the craft.

Let the velocity components due to self generated thrust be V_(C) i.e.V_(TAS) along longitudinal axes, V_(l) along lateral axis and V_(v)along vertical axis.

Similarly the wind vector will have spatial velocity components alongeach of the body fixed axis of the craft. Let V_(W) is the wind vectorin the vicinity of the air vehicle. V_(WH) is component of V_(W) onplane AB (horizontal plane) and V_(WV) is component of V_(W) on plane BC(vertical plane). β & γ are the angles made by spatial velocitycomponents of V_(W) with plane AB and plane BC respectively.

So, V _(WH) =V _(W) cos β

And V _(WV) =V _(W) cos γ

FIG. 4B.

V_(WH) makes an angle τ with OA (longitudinal axis). V_(WH) is resolvedinto two components. One component is along longitudinal axis. Its valueis V_(WH) cos τ. The other is along lateral axis of the air vehicle. Itsvalue is V_(WH) sin τ.

The sensor receiver along OA gives out put SV_(ƒ). It is vector sum ofTrue Airspeed V_(TAS) and component of V_(WH) along longitudinal axis.

SV_(ƒ) =V _(TAS) +V _(WH) cos τ  (24)

The sensor receiver along OB gives out put SV_(l). It is vector sum ofV_(l) and V_(WH) sin τ.

SV_(l) =V _(l) +V _(WH) sin τ  (25)

Similarly along OC,

SV_(v) =V _(v) +V _(WV) cos σ  (26)

As shown in FIG. 4B two sensor Rx Aes (27, 28), each placed at α°(clockwise) to SV_(ƒ) and SV_(l) sensor Rx Aes (23, 24) respectivelyhelp compute wind vector components along longitudinal axis and lateralaxis respectively. Let the outputs of these two sensors be S₁ and S₂respectively and they are in the same unit of measurement as thatspatial velocity of the craft.

The computations are done using the following equations.

Wind Vector Components:

For wind speed of V_(WH) Kms/Hour blowing towards a direction at τ°(clockwise) to aircraft heading, component along longitudinal axis,

V _(WH) cos τ=(S ₁ cos α−S ₂ sin α)−SV_(ƒ) cos 2α  (27)

And component along lateral axis,

V _(WH) sin τ=(S ₁ sin α+S ₂ cos α)−(SV_(l)+SV_(ƒ) sin 2α)  (28)

V _(TAS)=SV_(ƒ) −V _(WH) cos τ  (29)

Drift:

Eqn.-17 gives drift rate.

Wind Speed on Horizontal Plane

$\begin{matrix}{\left( V_{WH} \right) = \sqrt{\left( {V_{WH}\cos \; \tau} \right)^{2} + \left( {V_{WH}\sin \; \tau} \right)^{2}}} & (30) \\{\tau = {\tan^{- 1}\frac{V_{WH}\cos \; \tau}{V_{WH}\sin \; \tau}}} & (31)\end{matrix}$Direction of wind (Wind from direction)=(Aircraft heading w.r.t. truenorth+τ)+180°  (32)

If value exceeds 360°, then 360° is subtracted from the sum.

Course Correction for Drift Elimination:

Let, ‘a’ be the present heading of the craft and ‘w’ the wind towardsdirection. Then course correction required for maintaining the desiredtrack is

$\begin{matrix}{{{Course}\mspace{14mu} {correction}\mspace{14mu} {Angle}\text{:}\mspace{14mu} \Delta \; a} = {\sin^{- 1}\frac{V_{WH}{\sin \left( {w - a} \right)}}{V_{TAS}}}} & (33) \\{{{Course}\mspace{14mu} {to}\mspace{14mu} {steer}\text{:}\mspace{14mu} c} = {w - \left( {a + {\Delta \; a}} \right)}} & (34)\end{matrix}$

Similarly, other two additional receivers S₃ & S₄ with antennae alignedon vertical plane at an angle of α with respect to receiver antennae(28, 29) will help compute wind vector component on vertical plane andalong OC. Component along OC,

V _(WV) cos σ=(S ₃ cos α−S ₄ sin α)−SV_(v) cos 2α  (36)

and orthogonal to it on the game plane,

$\begin{matrix}{{V_{WV}\sin \; \sigma} = {\left( {{S_{3}\sin \; \alpha} + {S_{4}\cos \; \alpha}} \right) - \left( {{SV}_{l} + {{SV}_{v}\sin \; 2\; \alpha}} \right)}} & (37) \\{V_{WV} = \sqrt{\left( {V_{WV}\cos \; \sigma} \right)^{2} + \left( {V_{WV}\sin \; \sigma} \right)^{2}}} & (38) \\{\sigma = {\tan^{- 1}\frac{V_{WV}\cos \; \sigma}{V_{WV}\sin \; \sigma}}} & (39)\end{matrix}$

However, keeping economy, space and operational requirement in view thereceivers S₃ & S₄ may not be used. For zero pitch angle and wings levelcondition SV_(v) gives the wind component along OC and path anglecorrection for maintaining predefined height can be calculated fromcomputed angle of climb and descent (eqn.-21). These extra sensors arerequired for air, marine, sub-marine vehicles i.e. vehicles, movingthrough fluid medium and not for other types of vehicles.

Angle of Attack:

Angle of Attack is difference of Pitch Angle (Eqn.-11) and Angle ofClimb/Descent (Eqn.-21).

Track Angle Error (TAE):

It is the difference between desired track and actual track. Hence, itis the angle between bearing to next way point from current craftposition and present track heading.

Off Track Distance:

Direct Segment:

Angle between azimuth from next way point to current craft position andreverse direction of route segment is determined. Distance from currentcraft position to next way point is determined. Let, the angle be δ anddistance h. So, the perpendicular distance to the desired track i.e.,

Off Track Distance is p=h sin δ  (40)

During Turn:

Difference between Turn radius (eqns.-22 and 23) and actual distance toorigin of the turn computed from craft position coordinate andcoordinate of origin of turn radius gives the Off Track Distance duringturn.

Display 1. Avionics (FIG. 11)—

The conventional panel instruments are substituted by graphical display.The names are only suggestive of corresponding conventional panelinstrument function.

Failure Indication (101):

Gives out light flash to indicate a failure condition. Details offailure are enumerated under warning details (146)

Heading Indicator (102):

Aircraft true and magnetic heading are numerically displayed in themarked boxes.

Airspeed Indicator (103, 105 & 112):

TAS is numerically displayed (103). To the right of Attitude indicator(104) the True Airspeed (TAS) Bar (105) is displayed. A vertical Barwith a horizontal arrow on its head moves on the scale on which vitalspeeds for various status of flight such as stall, max endurance,maximum distance, normal and maximum limit speeds are marked. To theright of and attached to the TAS bar is a slender vertical bar thatindicates the prospective TAS in next five seconds and the prospectiveTAS (112) is displayed numerically below the TAS bar (105).

Attitude Indicator (104):

It is located at centre of Avionics Section. The aircraft icon is at thecentre of it. The graphical look is same as conventional AttitudeIndicator. Level of horizon is indicated by two straight lines each inline with either of the wings of the icon. The attitude of the aircraftis indicated by position of the aircraft icon for bank, climb anddescent with respect to horizon.

Wind Vector (106):

Instantaneous relative Wind direction and speed (107), Angle of Attack(108), wind vector components along longitudinal axis (109) and alonglateral axis (110) as well as drift rate (111) are numericallydisplayed.

Turn Indicator (122):

Turn rate in degrees/second is displayed numerically.

Bank Angle (123):

Accurate bank angle to left or right is displayed.

Turn Coordinator:

Three lights are placed below the attitude indicator to indicate whetherpath on the turn segment is actually followed or not. The centre light(126) gives on path indication. If deviating to inside of the turn thenthe light to turn side (125/127) comes on and when deviating outwards(opposite) to turn side the light opposite to turn side (125/127) comeson; the later being a sure indication for skid condition.

Vertical Information:

All information related to vertical movement and position are displayedbelow the Attitude Indicator (104). Angle of climb/descent (118),Vertical Speed (119), Altitude (121), terrain obstruction height ontrack heading (116), Clearance height over obstruction (115), verticalinclination indicator (113, 114) to indicate whether the craft is movingwith zero climb/descent angle or positive or negative angle indicated bythe position of the arrow with respect to the point. In the pathalignment indicator (120) an icon moves in proportion to vertical heightoffset from predetermined height planned and by moving controls to bringthe icon to cross ensures correct vertical path position for the craft.

2. Navigation Information: (FIG. 11)

Position Information:

Position of aircraft (128) in longitude and latitude, position of nextway point (129) in longitude and latitude and in terms of bearing anddistance from aircraft present position, position of alternate runway(130) (airfield) for the current route segment in bearing and distancefrom aircraft current position. Also, the current position is displayedin moving map display section (154)

Time Information (133):

Time related to take off airfield (134), last way point (135), next waypoint (136), 2^(nd) way point (137) are displayed numerically. The timesto be displayed are take off time (138), time elapsed since take off(139), time over last way point (140), time elapsed since crossing lastw/p (141), ETA over next w/p as per flight plan (142), present estimateof time over next w/p (143), speed adjustment required to meet ETA(planned) at next w/p (144), similar information for 2^(nd) next w/p(145, 146 & 147).

3. Control Guidance: (FIG. 11)

Horizontal Control Guidance (157):

The control guidance information is meant to provide guidance to thecrew for maintaining a predetermined track. The information provided aredirection of ground track (true and magnetic) (155), ground speed (157),Track Alignment Indicator TAI (159) which has an icon to indicatehorizontal displacement from track and the crew of the craft can takethe icon as aim figure to be kept on the cross to be on the centre lineof the track, Track Angle Error TAE (161), off track distance (162),heading correction/course to steer for maintaining track) (163). Theinformation windows (164) and (165) are also meant for control guidanceto the crew, hence placed under this group.

Turn Anticipation, Commence and Level Off Indicator (164):

trn flashing is meant to indicate aircraft is approaching way point toenter turn, com flashing indicates to commence turn and out flashingindicates to level off. The box is meant to provide turn direction(left/right) and roll out heading numerically.

Climb/Descent Anticipation, Commence and Level Out Indicator (165):

Either of ‘cl’ or ‘ds’ of cl/ds flashes to alert approach of w/p tocommence climb or descent, com flashing indicates to commence climb ordescent, and out flashes to indicate to complete manoeuvre. The box ismeant to provide level out altitude. Path Alignment Indicator PAI (120)in FIG. 10, is also a control guidance facility for maintainingpredetermined altitude by keeping the aim figure at the cross.

4. Additional Information and Aid:

Communication (155):

Route (main and stand by) communication frequency and ATC communicationfrequency and station identity are displayed.

Flight Plan, Check List & Other Reference Information (152):

Complete Flight Plan with ‘From-To’, way points, alternate runwaysinformation are displayed. The next w/p is highlighted. So is thealternate runway for the current route segment in which the craft isflying. It facilitates pre flight loading of flight plan as well asediting flight plan during flight without disturbing the normaloperation. The crew can select ‘check list’ and desired checklist asselected from a sub-menu will be displayed in the same window, otherreference information will be available on the same window uponselection.

Warning (151):

As and when a failure or otherwise warning condition arise the flashinglight draws attention of the crew and brief detail of warning isdisplayed.

-   (i) Drift (ii) Collision (iii) Weather (iv) Any warning from ATS (v)    Equipment failure etc. are conditions for warning.

Advice (150):

Advice received from Air Traffic Management (ATM) agency through datalink as well as those self generated are displayed here.

5. Navigation from Start to Destination:

Navigation and control guidance during Taxi out, Line up and departure(SID), Route and STAR (standard arrival) segments are facilitated bydisplay of aircraft icon at the centre of appropriate map displayed(154) depicting current aircraft position and movement of map/chart inproportion to aircraft displacement and through graphical display ofinformation as given under ‘control guidance’, position and timesection. FIG. 12 depicts display of Avionics, Navigation, Control andAdditional Information on a single screen.

Non Precision Approach:

Landing Guidance:

ABLAN System provides two types of non-precision landing guidancesimultaneously.

-   (i) Screen displayed guidance-   (ii) Interactive Voice Response (IVR) system audio guidance.

Each is self sufficient to provide non-precision landing support ofcategory I, II and III and is independent of the other. ABLAN System cancater for any temporary change in TDP position and usable length of therunway in use which can be set manually after receiving ATC advice orNOTAM.

Screen Displayed Guidance:

The normal subject centric moving map display (154) continues. Anextended centre line starting from R/W touchdown point coordinate witharrow head at TDP is displayed on moving map display (154) for referenceas track to be aligned with.

The approach and landing aid window opens to the right of the Moving MapNavigation window (154) in the horizontal control guidance window (157)with Runway heading up map display.

FIG. 13A.

It has display of (a) R/W icon (56) with marked touchdown point (58),(b) runway area landmarks and obstructions—and (c) an extended centreline (63) starting from R/W touchdown point.

Track Guidance:

All that the pilot has to do is to follow the extended center line (63)as track by aligning aircraft icon (64) with it. The Approach fixes,initial (62), intermediate (61), final (60) and missed approach point(59) are marked on the extended centre line.

Path Guidance:

FIG. 13A shows the icon ^(—)↑^(—) (57) generated at the point where theaircraft will touch the ground if present glide angle and rate ofdescent is maintained. The pilot has to adjust rate of descent so thatthe blank of the icon ^(—)↑^(—) (57) aligns with the TDP marker

(58) on track to give this view (

).

Position of craft on path is determined by following method. As shown inFIG. 13B T (66) is the present spatial position of the aircraft. Thecorresponding ground position is S (64). ST (65) is the altitude aboveR/W altitude (AMSL). TQ (68) is the present path on current trackheading, Q being the projected TDP (^(—)↑^(—)) (57).

$\begin{matrix}{{\tan \; \theta_{4}} = {{\frac{ST}{SQ}\left( {{\theta_{4}(67)} = {{angle}\mspace{14mu} {of}\mspace{14mu} {descent}}} \right)} = {{> {SQ}} = \frac{ST}{\tan \; \theta_{4}}}}} & (41)\end{matrix}$

Thus, position of Q is marked on the landing R/W map and at point Qprojected aim figure icon for path guidance (^(—)↑^(—)) (57) isdisplayed.

Audio Guidance

Track Guidance:

An Interactive Voice Response (IVR) system simulates ground radarcontroller. The following is the method for track guidance. Bearing ofthe TDP (58) from current position of the craft (64) is compared withrunway direction.

Found equal: voice on track.

Found not equal: Bearing of TDP is compared with (Runway heading+180)

Found in the domain [R/W direction & (R/W+180) direction]

Voice—Left of track.

Found not in the domain [R/W direction & (R/W+180) direction]

Voice—Right of track.

FIG. 14 A illustrates method for determining actual off track distance.

S (64) is the momentary aircraft ground position. P (58) is the TDP. SPis the ground distance to TDP. SN is the normal to PN (R/W extendedcentre line)<SPN=θ₅ (70) and

SN=SP sin θ₅ in metres  (42)

PN=SP cos θ₅  (43)

Voice announces the individual digits of the measured distance (SN).Sequence of voice guidance for track—ON/Left oft/Right of—Track, XYZMetres.

When PN value approaches Missed Approach Point Distance from TDP, IVRwarns the crew.

Path Guidance:

FIG. 14 B illustrates method for path guidance. T (66) is the spatialmomentary position of the aircraft. S (64) is the corresponding positionon ground. P (58) is the TDP.

ST is ht. of aircraft over ground found from the difference of aircraftaltitude (AMSL) and R/W altitude (AMSL)

SP is the ground distance to TDP. ST′ (73) is the desired ht. fordesired glide angle θ₆ (74):

$\begin{matrix}{{{ST}^{\prime} = {{SP}\; \tan \; \theta_{6}}}\begin{matrix}{{When},{{{ST} - {ST}^{\prime}} = 0}} & {{voice}\text{-}{on}\mspace{14mu} {{path}.}} \\{\neq {0\mspace{14mu} {and}\mspace{14mu} ( + ){ve}}} & {{voice}\text{-}{high}\mspace{14mu} {on}\mspace{14mu} {path}} \\{\neq {0\mspace{14mu} {and}\mspace{14mu} ( - ){ve}}} & {{voice}\text{-}{low}\mspace{14mu} {on}\mspace{14mu} {{path}.}}\end{matrix}} & (43)\end{matrix}$

|ST−ST′|→ht difference in metres.

Guidance sequence: high/low/on path. XYZ metres.

As craft approaches ST′=decision height, voice advices: Decision.

Abort (Missed Approach)

A Missed Approach is followed by increase in altitude and speed. This isinterpreted by the system and moving map display is changed to Area Map,control guidance window is restored and missed approach procedure isfollowed.

Holding Pattern

A point is selected in coordinates. A holding pattern figure isgenerated and displayed on the moving map for guidance of the crew.

Taxiing in

As the aircraft completes 100 meters run on the runway after TDP bothNavigation and Landing windows disappears from the screen and airfieldmap appears on the screen facilitating taxiing back to parking area.

Integration with AFCS (Auto Pilot):

Data and command generated will be fed to AFCS for control of aircraft.

Integration with Communication Capable of Digitalized Data Transfer:

Integration with ATN or any other suitable means for digitalized dataexchange will facilitate (i) exchange of position and movement databetween aircraft directly or through ground agency for ACAS type I andII service, (ii) Cockpit Display of Target Information (CDTI) functionto provide crew the current traffic environment, (iii) AutomaticDependence Surveillance (ADS) service to provide Air Traffic Management(ATM) and aircraft fleet owner the position and movement information inrespect of the aircraft.

Fail Operational System:

By operating two numbers of the system of present invention in tandem, afail operational configuration will be achieved.

CONCLUSION

Though the embodiment for use in aircraft has been described, the use isnot exclusive and does not exclude uses other than which have beendescribed. The sensor sub-system can be used in remaining types ofvehicles (surface, sub-surface, marine, sub-marine and space) fordetermining movement vector of the fluid (air and/or water) mediumthrough which the mount vehicle moves and 3-Dimensional spatial velocityof the mount vehicle and as such related movement, orientation, positiondata can be computed and start to destination navigation, control andguidance aid for vehicles operating on surface of the earth, in tunnels,on water surface, below water surface and space will be provided by thesystem as would be appreciated by the ones skilled in the art.

I claim:
 1. A method for self contained precision navigation and controlguidance from start to destination around the earth continuously forvehicles moving in air, in space, on surface of the earth, under surfaceof the earth (sub-surface), over water surface (marine) and under watersurface (sub-marine) comprising, Computation of 3-dimensional spatialvelocity of millimetre accuracy of the said vehicle along each of itsbody fixed axes. Determination of orientation (yaw, roll, pitch angles)of the said vehicles. Computation of movement data of fluid (air and/orwater) in the immediate vicinity of subject vehicle and effect of fluidmovement on the movement of the subject vehicle moving through the saidfluid(s). Derivation of Navigation Information in respect of the subjectvehicle for Navigation and Control guidance of the subject vehicle.Determination of momentary geophysical position on the earth's surfacein coordinates (longitude and latitude based on WGS-84 datum) & in termsof bearing and distance to a way point. Visual numeric or non-numericdisplay of navigation and position information for control and guidance.Display of geophysical momentary position on moving map for navigationand control guidance. Audio guidance during critical phase of navigationto enable the crew to be vigilant about the surrounding. Integrationwith AFCS (for control) & means for digitalized data communication (forCDTI, ADS & ACAS service) for air vehicles.
 2. A method according toclaim 1, wherein 3-dimensional spatial velocity of a vehicle along itsbody fixed axes will be computed by a self contained sensor sub-systemwhich will determine propagation time delay proportional to vehiclespeed, of a predetermined time frame pulse modulating EM wave carriertravelling from transmitting antenna to receiving antennae.
 3. A methodaccording to claim 2, wherein the sensor sub-system will measurepropagation time delay using Time to Digital Converters (TDC), the saiddelay being a function of spatial velocity of the mount vehicle alongrespective body fixed axis.
 4. A method as in claim 2, wherein thecomputed time delay for each of the axis divided by duration of pulse(one cycle) of reference frequency will directly read spatial velocityof the vehicle along its respective body fixed axes.
 5. A method as inclaim 2, wherein the sensor sub-system will comprise of a second set oftransmitter and receiver that will operate with a time frame pulse ofdifferent duration for self correction of position information to annuldrift integration of positional error (along track, cross track andvertical).
 6. A method according to claim 2, wherein the sensorsub-system will consist of two sets of sensor. Each set will have singlevery low power transmitter feeding single/plurality of TransmitterAntennae and plurality of low sensitivity receivers having antennaealigned with body fixed axes or lines parallel to body fixed axes of thevehicle and that the said transmitter, receivers, their antennae andother sensor hardware will be placed inside a metal casing to make thesystem a zero emission to space and non-susceptible to electronicjamming system.
 7. A method according to claim 1, wherein theorientation (pitch, roll yaw angles) of the vehicle with respect togeo-physical plane will be derived by using 3-dimensional spatialvelocity along body fixed axes of the vehicle and well knownmathematical formulae.
 8. A method according to claim 1, whereindirection and speed of fluid (air and/or water) movement in theimmediate vicinity of mount vehicle affecting movement of the mountvehicle will be determined using two additional receivers for horizontalplane and two for vertical plane, the antennae of which will be placedat a predetermined offset angle with respect to receiver antennae meantfor sensing spatial velocity along longitudinal, lateral and verticalaxes. The data provided by this pair of additional receivers will helpto compute (a) fluid (air and/or water) vector (b) components of fluidvector along longitudinal, lateral and vertical axes of the vehicle (c)the drift rate, (d) relative velocity and movement of the vehicle withrespect to fluid e.g., TAS and angle of attack for air vehicle (e)course correction required to maintain desired track and (f) courseheading required to steer for maintaining desired track.
 9. A method asin claim 1, wherein Navigation information will be determined which inrespect of vehicles of airborne category are i) magnetic and trueheading ii) course on ground (both magnetic and true) iii) True AirSpeed iv) ground speed v) altitude vi) angle of climb/descent vii)vertical speed viii) angle of bank ix) turn direction x) turn rate xi)attitude and xii) momentary geo-physical position, using sensor data andcommercial computer encompassing real time operating system and highlevel language to solve mathematical problems.
 10. A method as in claim9, wherein the heading, track, altitude and geophysical position aredetermined by dead reckoning (DR) from a position initialized at startup or line up point and reference to external system or instrument willnot be required for correction of these information during the entireoperation period that follows.
 11. A method as in claim 1, whereinnavigation information (mentioned in claims 7, 8 and 9) will bedisplayed graphically as substitute for multiple primary and additionalflight data instruments on display screen.
 12. A method as in claim 1,wherein Navigation and control guidance will be provided under alloperating conditions to the subject aircraft from start to destinationaround the earth by display of (i) path overlay and way point on movingmap display wherein aircraft position is marked by an icon (ii) positioninformation in coordinates (longitude and latitude based on WGS 84datum) in respect of subject vehicle and relevant way points (iii)bearing and slant/ground distance to any way point from the craft orfrom another way point iv) off track distance (v) track angle error and(vi) prospective manoeuvre indication.
 13. A method as in claim 12,wherein prospective manoeuvre indication includes (i) course correctionto counter drift (ii) turn and climb/descent anticipation, guidance forturn and climb/descent from commencement to completion of turn andclimb/descent and (iii) for hold pattern.
 14. A method as in claim 13,position of origin of turn radius for ‘fly by’ and ‘fly over’ turn isdetermined for ATS route designator published turn radius as well as forother instances and continuous computation of distance to origin of turnis carried out to facilitate a controlled turn for the air vehiclewithin specified envelope.
 15. A method as in claim 1, wherein toprovide non-precision approach and landing guidance of Cat I, II & IIIstandard for any runway around the globe for which longitude andlatitude information are available by display of extended runwaycenterline for track alignment and an aim figure to be aligned withmarked runway Touch Down Point (TDP) for path guidance.
 16. A method asin claim 1, wherein to provide virtual ground controller controlledapproach guidance of Precision Approach Radar (PAR) type for any runwayaround the globe for which Coordinate (Long., Lat.) data is available byproviding continuous audio guidance during landing approach for trackand glide slope alignment along with information regarding horizontaland vertical displacement from path and track through Interactive VoiceResponse (IVR) system.
 17. A method as in claim 1, wherein position andmovement data will be exchanged with traffic elements in the surroundingand transmitted to ground controlling agencies using digitalized datatransfer through suitable communication means to facilitate display oftarget position on moving map section of display unit (CDTI), to computethreat to subject air vehicle from traffic element (ACAS) and tofacilitate Automatic Dependent Surveillance (ADS) service.
 18. A methodas in claim 1, wherein derived navigation information will be fed toAutomatic Flight Control System (AFCS) for automatic control of airvehicle.
 19. A method as in claim 1, wherein flight plan can be made forpoint to point navigation having facility for selecting multiple numberof way points and alternate field(s) for a particular route segment andedit change to flight plan without affecting normal flight operation andeffect changes when so desired.
 20. A method as in claim 1, wherein thesystem can be used to provide navigation and control aid to surface(rail and road), marine, sub-marine or space vehicles too.